Showing papers on "Wing root published in 2008"

Aircraft Control via Variable Cant-Angle Winglets

Abstract: This paper investigates a novel method for the control of “morphing” aircraft. The concept consists of a pair of winglets with adjustable cant angle, independently actuated and mounted at the tips of a baseline flying wing. The generalphilosophybehindtheconceptwasthatforspecific flightconditionssuchasacoordinatedturn,theuseoftwo control devices would be sufficient for adequate control. Computations with a vortex lattice model and subsequent wind-tunnel tests demonstrate the viability of the concept, with individual and/or dual winglet deflection producing multi-axis coupled control moments. Comparisons between the experimental and computational results showed reasonable to good agreement, with the major discrepancies thought to be due to wind-tunnel model aeroelastic effects.

Influence of Finite Span and Sweep on Active Flow Control Efficacy

Abstract: Active flow control efficacy was investigated by means of leading-edge and flap-shoulder zero mass-flux blowing slots on a semispan wing model that was tested in unswept (standard) and swept configurations. On the standard configuration, stall commenced inboard, but with sweep the wing stalled initially near the tip. On both configurations, leading-edge perturbations increased CL,max and post stall lift, both with and without deflected flaps. Without sweep, the effect of control was approximately uniform across the wing span but remained effective to high angles of attack near the tip; when sweep was introduced a significant effect was noted inboard, but this effect degraded along the span and produced virtually no meaningful lift enhancement near the tip, irrespective of the tip configuration. In the former case, control strengthened the wingtip vortex; in the latter case, a simple semi-empirical model, based on the trajectory or "streamline" of the evolving perturbation, served to explain the observations. In the absence of sweep, control on finite-span flaps did not differ significantly from their nominally twodimensional counterpart. Control from the flap produced expected lift enhancement and CL,max improvements in the absence of sweep, but these improvements degraded with the introduction of sweep.

Mesh Generation and Deformation Algorithm for Aeroelasticity Simulations

Abstract: H IGH-FIDELITY flow solvers for aeroelastic applications require the use of computational meshes that deform as the structure is being displaced. High-aspect-ratio wings increase the demands on the robustness of the mesh-deforming algorithm, because these wings are extremely flexible and attain deformations that are a significant fraction of the span of the wing. The mesh deformation algorithm must be not only robust but also computationally inexpensive to avoid penalizing the turnaround time of the aeroelastic computations. Different approaches have been developed to solve the movingmesh problem. For meshes generated by using overlapping grids, a natural way to allow for gridmotion is to slide the overlapping region of the grids [1,2]. The advantage of thismethod is that the body-fitted meshes do not deform during the body motion. A disadvantage of this approach is that the interpolation algorithm that communicates the solution between grids has to be updated for each overlapping position of the grids. The tension spring analogy [3] is one of themostwidely usedmesh deformation strategies. In this approach, each edge of the mesh is represented by a spring for which the stiffness is proportional to the reciprocal of the length of the edge. By replacing the edges with springs, a deformation of the boundary translates into a deformation of the spring network, which adjusts its shape to the equilibrium position of the network. The displacements in each direction are decoupled and the equation that updates the position of the nodes is relatively easy to solve. A disadvantage of this method is that for highly distorted meshes, collapsed or negative volume cells may appear, especially on high-aspect-ratio cells such as those used for viscous flows. An improvement over the tension spring analogy is the torsion spring analogy [4,5]. The torsion spring analogy consists of adding a torsional spring to the tension-spring-analogy technique. The stiffness of the torsional spring is related to the angle between the edges. As the angle tends to zero, the stiffness tends to infinity, thus preventing vertices from crossing over edges and avoiding negativevolume cells. The disadvantage of this method is the higher complexity and computational cost than with the tension spring analogy. The transfinite-interpolation mesh deformation technique is based on the linear interpolation of the boundary motion [6]. The motion of a node located between amoving and a fixed boundary is equal to the motion of themoving boundary times a scale factor. This scale factor, assigned to each node of themesh, depends on the distances from the node to the moving and the fixed surfaces. The scale factor is 1.0 for nodes on the moving boundary and 0.0 for nodes on the fixed boundary. The method guarantees a smooth transition between the moving boundaries and the fixed boundaries. One disadvantage of this method is that it cannot guarantee the mesh orthogonality at deforming surfaces, a condition that is important for viscous flows. Another approach to simulatemesh deformation is to use the linear elasticity equations [7]. The deformed grid is obtained by solving the equilibrium equations for the stress field. Themodulus of elasticity is chosen to be inversely proportional to the cell volume or to the distance from the deforming boundaries. Therefore, the cells close to the moving boundaries have small deformations, and the majority of the mesh deformation is relegated to the regions farther away from the moving boundary. This Note presents a grid generation and deformation algorithm for wings with large deformations. The computational domain was discretized using a hybrid grid that consisted of structured hexahedra around the wing and unstructured triangular prisms elsewhere. The mesh was divided in layers that were topologically identical in the spanwise direction. The mesh deformation algorithm was applied in two steps. First, the spring analogy technique was applied to deform the nodes within a mesh layer. Second, the layers were deformed to be perpendicular to the boundaries of the domain and to the surface of the wing. The Note describes the mesh generation algorithm and the mesh deformation algorithm and shows results for a wing with large tip deformation.

Transonic Limit Cycle Flutter of High-Aspect-Ratio Swept Wings

Abstract: This paper is motivated by the observed transonic limit cycle flutter of a high-aspect-ratio swept wing tested at DLR, German Aerospace Center in Gottingen. We show that the aeroelastic mechanism responsible for the limit cycle flutter behavior can be explained in terms of the structural washout effect, which is strongly stabilizing at transonic Mach numbers. This limits the energy transfer rate to the wing, resulting in a limit cycle flutter mode in which bending and torsion are almost perfectly in phase, and the streamwise wing-section motion resembles a single-degree-of-freedom torsion mode with an axis of rotation forward of the leading edge. For swept wings, the apparent pitching motion of the wing sections arises naturally through the structural washout and is present in the first bending mode that enters limit cycle flutter. The aeroelastic washout effect may produce counterintuitive results, and increasing the dynamic pressure may actually be stabilizing. In such a case, decreasing the dynamic pressure would increase the limit cycle amplitude, and high-altitude transonic flutter becomes a real possibility. The results demonstrate the importance of using a nonlinear structural model in these calculations. If a linear structural model is used, the critical dynamic pressure at limit cycle flutter onset is overpredicted by a factor of nearly 3, and the predicted flutter mode is at a frequency much higher than was observed during wind-tunnel tests.

Topology Optimization Approach for the Determination of the Multiple-Configuration Morphing Wing Structure

Abstract: The paper introduces an innovative topology optimization approach for determining the distribution of structural properties and actuators to design amorphing wing that is capable of achievingmultiple target shapes. The previous investigation by the authors demonstrated, using various problem formulations and a novel modeling concept, the fundamental topology synthesis of a simple two-configurationmorphingwing structure.Theprimary objective of the present investigation is therefore to introduce improvements and extensions to the previous concepts and problem formulations to those capable of accommodating the multiple-configuration definitions. The investigation includes the formulation of appropriate topology optimization problems and the development of effective modeling concepts. In addition, principal issues on the external load dependency and the reversibility of a design, as well as the appropriate selection of a reference configuration, are addressed in the investigation. The methodology to control actuator distributions and concentrations is also discussed. Finally, an examplemultiple-configuration problem that portrays the generic surveillance mission is solved to demonstrate the potential capabilities of the approach.

Calculated Viscous and Scale Effects on Transonic Aeroelasticity

Abstract: A viscous-inviscid interactive coupling method is used for the computation of unsteady transonic flows. A lagextrainment integral boundary layer method is used with a transonic small disturbance potential code to compute the transonic aeroelastic response for two wing flutter models. By varying the modeled length scale, viscous effects may be studied as the Reynolds number per reference chordlength varies. Appropriate variation of modeled frequencies and generalized masses then allows comparison of responses for varying scales or Reynolds number. Two wing planforms are studied: one a four percent thick swept wing and the other a typical business jet wing. Calculations for both wings show limit cycle oscillations at transonic speeds in the vicinity of minimum flutter speed indices.

Nonlinear Response Structural Optimization of a Joined Wing Using Equivalent Loads

Abstract: The joined wing is a new concept of the airplane wing. The forewing and the aft wing are joined together in the joined wing. The joined wing can lead to increased aerodynamic performances and reduction of the structural weight. The structural behavior of the joined wing has a high geometric nonlinearity according to the external loads. Therefore, the nonlinear behavior should be considered in the optimization of the joined wing. It is well known that conventional nonlinear response optimization is extremely expensive; thus, the conventional method is almost impossible to use for large-scale structures such as the joined wing. In this research, geometric nonlinear response optimization of a joined wing is carried out by using equivalent loads. The used structure is a joined wing that is currently being developed in the U.S. Air Force Research Laboratory. Equivalent loads are the load sets that generate the same response field in linear analysis as that from nonlinear analysis. In the equivalent loads method, the external loads are transformed to the equivalent loads for linear static analysis, and linear response optimization is carried out based on the equivalent loads. The design is updated by the results of linear response optimization. Nonlinear analysis is carried out again and the process proceeds in a cyclic manner until the convergence criteria are satisfied. It was verified that the equivalent loads method is equivalent to a gradient-based method; therefore, the solution is the same as that of exact nonlinear response optimization. The fully stressed design method is also used for nonlinear response optimization of a joined wing. The results from the fully stressed design and the equivalent loads method are compared.

Transition Mechanisms for Transport Aircraft

Abstract: This paper outlines basic transition mechanisms encountered on transport aircraft. Transition is highly dependent on operating conditions, wing and airfoil geometry, and surface conditions, and any prediction scheme must accurately account for the relevant physics. Moreover, the efficacy of control depends on the physics of the transition process– whether one is delaying transition through LFC or enhancing turbulence for propulsion systems or separation control. Four basic instability mechanisms generally lead to transition on a swept transport-aircraft wing: Tollmien-Schlichting, crossflow, attachment line, and Gortler instabilities. Other drivers include juncture flows (for example at the wing root), locally supersonic flow on the suction side of the wing, and surface features.

Folding Wing Root Mechanism

Abstract: Improvements to the hinge of a folding aircraft wing including a load bearing hinge mechanism with multiple locking mechanisms. The mechanism includes rigid panels that cover the hinge area when the wing is deployed, and the wing itself covers the hinge area when the wing is retracted. A control lever is used to actuate the wing and incorporates safety features to prevent unwanted actuation.

Dynamics of Composite Aircraft Wings Carrying External Stores

Abstract: This paper focuses on the free vibration and dynamic response to external time-dependent loads of aircraft wings modeled as thin-walled anisotropic composite beams carrying eccentrically located heavy stores. In this context, bending-twisting coupling induced by both the eccentric heavy stores arbitrarily located along the wing span and chord and by the anisotropy of the constituent material that is essential when dealing with aircraft wing problems was included in the approach of the problem. In addition to the anisotropy of constituent materials, transverse shear and warping restraint effects were also incorporated. The governing equations of the wing-store system and the related boundary conditions are derived via Hamilton's principle. To solve the eigenvalue/boundary problems, the extended Galerkin method is applied. Numerical simulations highlighting the implications of external stores coupled with the implementation of the structural tailoring on eigenfrequency and dynamic response to external time-dependent loads are supplied and pertinent conclusions are outlined.

Insect-inspired wing actuation structures based on ring-type resonators

Abstract: In this paper, we illustrate and study the opportunities of resonant ring type structures as wing actuation mechanisms for a flapping wing Micro Air Vehicle (MAV). Various design alternatives are presented and studied based on computational and physical models. Insects provide an excellent source of inspiration for the development of the wing actuation mechanisms for flapping wing MAVs. The insect thorax is a structure which in essence provides a mechanism to couple the wing muscles to the wings while offering weight reduction through application of resonance, using tailored elasticity. The resonant properties of the thorax are a very effective way to reducing the power expenditure of wing movement. The wing movement itself is fairly complex and is guided by a set of control muscles and thoracic structures which are present in proximity of the wing root. The development of flapping wing MAVs requires a move away from classical structures and actuators. The use of gears and rotational electric motors is hard to justify at the small scale. Resonant structures provide a large design freedom whilst also providing various options for actuation. The move away from deterministic mechanisms offers possibilities for mass reduction.

Reduced Span Wings with Wing Tip Devices, and Associated Systems and Methods

Abstract: Reduced span wings with wing tip devices, and associated systems and methods are disclosed. A method for designing a wing in accordance with a particular embodiment includes establishing a target lift value for a winglet to be attached to a wing, the wing having a wing root, a wing tip and a twist distribution that results in a loading at the wing tip that is higher than a target loading level. The method can further include selecting a planform shape of the winglet to produce less of an increase in loading at the wing tip compared to the wing tip loading produced by other winglet planform shapes having the same target lift value. Accordingly, in particular embodiments, the winglet design can offset or at least partially offset the reduced wing tip loading produced by the twist distribution.

Aeroservoelastic Testing of a Sidewall Mounted Free Flying Wind-Tunnel Model

Abstract: A team comprised of the Air Force Research Laboratory (AFRL), Northrop Grumman, Lockheed Martin, and the NASA Langley Research Center conducted three j wind-tunnel tests in the Transonic Dynamics Tunnel to demonstrate active control technologies relevant to large, exible vehicles. In the rst of these three tests, a semispan, aeroelastically scaled, wind-tunnel model of a ying wing SensorCraft vehi- cle was mounted to a force balance to demonstrate gust load alleviation. In the second and third tests, the same wing was mated to a new, multi-degree-of-freedom, sidewall mount. This mount allowed the half-span model to translate vertically and pitch at the wing root, allowing better simulation of the full span vehicle's rigid-body modes. Gust Load Alleviation (GLA) and Body Freedom Flutter (BFF) suppression were successfully demonstrated. The rigid body degrees-of-freedom required that the model be own in the wind tunnel using an active control system. This risky mode of testing necessitated that a model arrestment system be integrated into the new mount. The safe and successful completion of these free-flying tests required the development and integration of custom hardware and software. This paper describes the many systems, software, and procedures that were developed as part of this effort.

Increasing the Lift-Drag Ratio of an Unmanned Aerial Vehicle Using Local Twist

Abstract: Recently published works predict that any planform shape may be optimized with twist to reduce the induced drag to an optimum value. When the twist is applied along the span of the airplane, the lift-drag ratio is lower than that with no twist. This can be corrected if twist is applied only in a specific portion of the span. The objective of this paper is to demonstrate that the local twist increases the lift-drag ratio using two different inviscid computational fluid dynamics codes and to describe the method employed to obtain the twist start line to increase the lift-drag ratio. The method was applied to an unmanned aerial vehicle designed for the early detection of oil leakages in the extraction areas, and a variation of 8 cm in the wing tip was obtained. The results show that the lift-drag ratio of the twisted wing is higher than that with no twist in conditions close to cruise flight. The lift-drag ratio increased 2.89 and 0.31%, estimated by Multhopp's method and by the vortex-lattice method, respectively. The results demonstrate that the local twist may increase the lift-drag ratio when it is applied in the way explained in the present paper.

Insect-inspired wing actuation structures based on ring-type resonators

Abstract: In this paper, we illustrate and study the opportunities of resonant ring type structures as wing actuation mechanisms for a flapping wing Micro Air Vehicle (MAV). Various design alternatives are presented and studied based on computational and physical models. Insects provide an excellent source of inspiration for the development of the wing actuation mechanisms for flapping wing MAVs. The insect thorax is a structure which in essence provides a mechanism to couple the wing muscles to the wings while offering weight reduction through application of resonance, using tailored elasticity. The resonant properties of the thorax are a very effective way to reducing the power expenditure of wing movement. The wing movement itself is fairly complex and is guided by a set of control muscles and thoracic structures which are present in proximity of the wing root. The development of flapping wing MAVs requires a move away from classical structures and actuators. The use of gears and rotational electric motors is hard to justify at the small scale. Resonant structures provide a large design freedom whilst also providing various options for actuation. The move away from deterministic mechanisms offers possibilities for mass reduction.

Examining Effects of Increased Effective Area on Performance of Movable Tip Strakes

Abstract: Nomenclature  = b=S, aspect ratio b = span, mm CD = wing drag coefficient CL = wing lift coefficient c = chord, mm d = angle relative to wing chord plane, deg q = 1=2 1V 1, freestream dynamic pressure, kPa Re = 1V1c= 1, Reynolds number based on wing model chord S = reference area, mm V1 = freestream velocity, m=s x, y, z = aerodynamic axes = angle of attack, deg = strake leading-edge sweep angle, deg = absolute viscosity, N s=m = air density, kg=m

Conceptual Morphing Aircraft Sizing Using a Multi-Level Optimization Strategy

Abstract: *† This paper presents an approach for sizing (i.e. predicting the dimensions, weight and performance) of a morphing aircraft based upon a multi-level design optimization approach. For this effort, a morphing wing is one whose planform can make significant shape changes in flight – increasing wing area by 50% or more from the lowest possible area, changing sweep 30o or more, and / or increasing aspect ratio by as much as 200% from the lowest possible value. The top-level optimization problem seeks to minimize the gross weight of the aircraft by determining a set of “baseline” variables – these are common aircraft sizing variables (T/W, S, AR, t/c, Λ, λ), along with a set of “morphing limit” variables – these describe the maximum shape change for a particular morphing strategy (e.g. Δb, Δc, ΔΛ). The sub-level optimization problems represent each leg in the morphing aircraft’s design mission; here, each sub-level optimizer minimizes fuel consumed during each mission leg by changing the wing planform within the bounds set by the baseline and morphing limit variables from the top-level problem. This approach predicts the wing empty weight based on a parametric equation derived from structural optimization studies of morphing wing concepts. This multi-level optimization sizing approach then sizes a morphing aircraft whose wing can change both sweep and wing root chord; this demonstration highlights the improved effectiveness of this multi-level approach over previous morphing aircraft sizing approaches.

Maneuver Load Control for Flexible Delta Wing using Multidisciplinary Design Optimization

Abstract: This paper addresses the Maneuver Load Control (MLC) for a generic fighter aircraft using optimization techniques to augment its performance. Multidisciplinary Design Optimization (MDO), an important application of which is in the field of aircraft design, is applied to the Maneuver Load Alleviation of a generic fighter aircraft, in order to reduce the wing loading and also to improve the performance. The aircraft considered is a typical generic multi role combat aircraft, with a delta wing and no horizontal tail. It is assumed that there are no individual control surfaces for pitch (elevator) and roll (aileron) control. Instead, both these controls are achieved by having two independent control surfaces viz. Inboard and Outboard ELEVONS (ELEVator+ailerON) on each wing. An optimized combination of differential elevon deflections are generated, which, in turn, would redistribute the loading on the wing so that that the wing root bending moment is reduced. The optimization problem has been formulated for (i) minimizing the wing root bending moment (ii) improving the flight envelope, by considering the angle of attack and the differential elevon deflections as design variables, subject to equality and inequality constraints. These constraints would arise from various interdisciplinary areas such as Aerodynamics, Structures and Controls. In the optimization cycle, for an initial guess, a global search is done using Genetic Algorithm (GA) and with this as initial value, the actual optimization is performed using Gradient Based Technique coupled with Method of Feasible Directions (MFD) for constraint handling, to generate the final optimized solution. The static aero elastic analysis is done using NASTRAN/PATRAN, while aerodynamic analysis is done using a Cartesian grid based EULER solver. The load computations are carried out for a few identified maneuvers of the aircraft, considering the inertial parameters such as vertical acceleration, roll rate and roll acceleration. The maximum bending and torque moment values on the wing are computed for various maneuver load cases and the critical maneuver loads are identified as corner points of the flight envelope boundaries. A few of these critical maneuver load cases have been identified for the present study and the MLC benefits achieved - structural as well as performance are presented. Further aerodynamic analysis is done for the optimized geometry for the cases studied, to compute the centres of pressure of the optimized solutions to confirm the reduction in the bending moments.

Aerodynamic Design of Three-dimensional Low Wave-drag Biplanes Using Inverse Problem Method

Abstract: The Busemann biplane is a representative airfoil which has possibility of realizing lowboom and low-drag. Aerodynamic designs based on the Busemann biplane are demanded for future supersonic transports. In this paper, extension of the Busemann biplane airfoil to 3-D wing is firstly investigated. Then an inverse problem method is applied to designing a high L/D biplane wing after the evaluation of a capability of the method for designing a practical 3-D supersonic biplane wing. Analyses and designs are based on Computational Fluid Dynamics (CFD). Free stream Mach number is 1.7. The authors achieved the following performance. At 0.11 of CL, L/D of a tapered wing combined with rectangular root region having span sections of Busemann biplane is 16.9, and L/D of the inversely designed biplane is 19.7. In order to improve three dimensional demerits about aerodynamic performance such as Mach cones, we attempted whether a 3-D biplane wing, the upper element of which was a flat plate wing, converged at a known target wing. Finally, L/D of the biplane wing was improved to 20.0 by setting up target pressure distributions aiming at a reduction of CD due to Mach cones around the wing root or kink.

Wing-fuselage section of an aircraft

Abstract: A wing-fuselage section of an aircraft, which wing-fuselage section comprises a wing root at which the wing of the aircraft is connected to the fuselage, a fuselage region with fuselage frame elements that extent across the longitudinal direction of the aircraft, and a wing region with spars that extend in the direction of the wingspan. According to the invention, the spars of the wing region and the fuselage frame elements of the fuselage region form part of an integral assembly that extends at least over a middle part of the wing and the fuselage region, including the wing roots.

Design Space Exploration of Supersonic Formation Flying Focusing on Drag Minimization

Abstract: This study investigates the drag characteristics of a supersonic formation-flying concept that aims to reduce wave drag and sonic boom via shock-wave and expansion-fan interaction. Because of the complex interaction patterns seen in three-aircraft formations, optimization was applied as a rational means for design. To consider both the cruise efficiency and the safety of the aircraft, the objective functions are chosen to be the total LID of the formation and the minimum separation distance among the aircraft. The design variables define the relative positions of the aircraft and, as for the coordinate definition, the skewed cylindrical coordinate system that has been proven to be very effective in extracting the physics of the flowfield was chosen. Optimization results show a good correlation with results from previous studies. However, optimization arrived at a formation that unexpectedly achieved high performances in regions of the design space in which the previous study suggested bad performance. These solutions took advantage of the design space of three-aircraft formations (i.e., the synergistic effects of the flowfield of the two leading aircraft) and exploited the difference between the design space of two-aircraft formations.

Pulping knife used for soya-bean milk machine

Abstract: A mincing knife for the soymilk machine comprises a wing root, the center of which is provided with a shaft hole, and a knife wing comprising the first knife wing, the second knife wring, the third knife wing and the fourth knife wing, and the cutting surface rotating around the knife wing is provided with an arc cutting edge, and the edge of the wing root between every two adjacent knife wings is provided with a cutter groove The first knife wing and the third knife fold up relative to the wing root plane, and the angle Alpha between the first knife wing and the wing root plane is between five degree and twenty degree, and the angle Beta between the third knife wing and the wing root plane is between ten degree and forty degree The second knife wing and the fourth knife wing fold down relative to the wing root plane, and the angle Gamma between the second knife wing and the wing root plane is ten degree and forty degree, and the angle Delta between the fourth knife wing and the wing root plane is between five degree and twenty degree The first knife wing and the third knife wing of the utility model fold upward relative to the wing root plane, and the second knife wing and the fourth knife wing fold downward relative to the wing root plane, thus to enlarge the length of the blade grinding working surface, increase the grinding efficiency, reduce the frictional resistance of the fluid, shorten the grinding time and reduce the power dissipation, thus to prolong the service life of the mincing knife

for Optimization Studies

Abstract: A new benchmark wing model for optimization algorithm comparisons that may include flutter and divergence, aeroelastic tailoring, buckling and post buckling, vibration and natural frequency analyses is proposed. The idea behind this wing model is that the laminate design of the wing root joint is used for the entire wing structural box (upper and lower skin panels, ribs, and front and rear spars). The structure is purposely over designed to provide a starting point for subsequent structural optimization. The wing model will be made freely available to researchers.

Transient Structure on Unmanned Combat Air Vehicle: Effect of Pitch-Up Motion

Abstract: F LIGHT envelopes of unmanned combat air vehicles (UCAV) planforms demand high performance at high angle of attack when undergoing dynamic motions. Up to this point, studies of the present UCAV planform have focused primarily on the flow structure on stationary planforms at defined angles of attack (static conditions), with varying Reynolds number. The dynamic effects of a perturbed or pitching motion on the flow structure of the planform have received less attention. Nearly all investigations have addressed the flow structure on a delta wing under dynamic conditions, whereby the wing has a relatively large sweep angle, in comparison with the low values of sweep of a UCAV planform. Among these, LeMay et al. [1] studied the dynamics of vortices on a pitching deltawing, andMiller andGile [2] examined the effect of blowing on delta wing vortices during pitching. A basic finding is that the location of vortex breakdown, compared with the static case, is further upstream during pitch-up motion and further downstream during pitch-down motion. This phase lag becomes larger with increasing pitch rate. Myose et al. [3] studied the effects of more complex planforms, including diamond, cropped, delta, and double-delta configurations. Emphasis was on vortex breakdown during pitching. Their main finding was that the cropped delta wing, which resembles the fuselage-apex form of the present UCAV, had the latest onset of leading-edge vortex breakdownduring pitching,whereas the doubledelta wing had the earliest onset. This observation may be one of the reasons underlying new designs of complex UCAV planforms. For a double-delta wing undergoing pitching motion, Ericsson [4] described the consequences of large-amplitude pitching oscillations at high angle of attack, whereas Grismer and Nelson [5] studied the aerodynamics of pitching motion with and without sideslip. Grismer et al. [6] and Hebbar et al. [7] also examined the influence of sideslip on the control of vortical flow structure on such planforms. Modern military aircraft involve complex, novel shapes such as the presentUCAV(approximating theBoeingX-45) to incorporate stealth technology. For such planforms, Elkhoury andRockwell [8] provided dye visualization images over a range of angle of attack and Reynolds number, accounting for the potential interaction of the coexistence of vortices emanating from the apex and the wing root, and the onset of vortex breakdown. Elkhoury et al. [9] employed a particle image velocimetry (PIV) technique to investigate the crossflow plane and near-surface topology by determining the mean and unsteady representations of their flow structure at various ranges of angle of attack and Reynolds number. However, none of the investigations so far have addressed the influence of unsteady maneuvers, such as the effect of a sudden increase of the angle of attack on the flow structure. Hence, the aimof the present investigation is to provide insight into the transient flow structure at a specific location normal to the planform surface, which takes place on a generic aerodynamic planform undergoing pitch-up motion at various ramp rates.

Design and Fabrication of Dragonfly Test Bed for Aerodynamic Characterization

Abstract: This paper focuses on the design of an electro-mechanical device for studying the aerodynamic behavior of flapping wings. The experimental device is designed to mimic the flight behavior of dragon fly. Wing flapping speed is precisely controlled by controlling the motor speed. Wing flapping amplitude could be varied by changing the rotating arm length. Wing rotation amplitudes during downand upstroke could be different and are controlled separately by two different springs. A six degree of freedom sensor is placed at the wing root to collect the force and torque data. The test of a wing with a dragonfly hind-wing contour but enlarged 11 times, showed the device met the design expectation, and further more, the phase-averaged data for lift force in one flapping cycle had the similar pattern as the ones obtained via CFD simulations as well as the one calculated based on a real dragonfly’s flight behavior.

Surface pressure measurement on a wing of SWIM by using PSP

Abstract: this study, three dimensional surface pressure distributions of SWIM whose main wing has NACA4412 airfoil with NACA0012 flaps were experimentally measured by pressure sensitive paint. Surface pressures on suction and pressure sides of the wing were measured by changing an angle of attack at a Reynolds number of 3.1x105 in KARI 1m subsonic wind tunnel. The experimental results showed that as an angle of attack increases minimum pressure region on a suction side moved from the wing root to the tip and low pressure region around trailing edge of the wing tip which causes wing tip vortex was observed. Although low pressure region at the tip still observed at an angle of attack 15 deg., other area on a suction side showed flat pressure distribution in a span-wise direction. It was also observed that the mean value of pressure coefficients was about 0.077 through a comparison between PSP and pressure taps at the same test conditions.

Sail wing with high span efficiency and controlled pitching moment

Abstract: An aircraft comprising a fuselage, a sail wing appended to the fuselage, the sail wing having a sail wing root chord length, and wherein the sail wing includes a sail wing leading edge spar, a sail wing membrane attached to the sail wing leading edge spar, and a sail wing trailing edge wire located at a trailing edge of the sail wing membrane, the aircraft further comprising a wing surface extension, located aft and at an inboard area of the sail wing trailing edge wire, the wing surface extension having a wing surface extension root chord length, and wherein the wing surface extension includes a wing surface extension membrane attached to the sail wing trailing edge wire, and a wing surface extension trailing edge, and wherein the wing surface extension trailing edge is reflexed such that the wing surface extension trailing edge is positioned upwards at a first angle with respect to a plane formed along a centerline of the aircraft and along the lower surfaces of the sail wing. The wing surface extension is further configured to affect a reduced nose-up pitching moment, and to produce a more even coefficient of lift along the wing of the aircraft.

Subsonic transport composite wing model for optimization studies

Abstract: A new benchmark wing model for optimization algorithm comparisons that may include flutter and divergence, aeroelastic tailoring, buckling and post buckling, vibration and natural frequency analyses is proposed. The idea behind this wing model is that the laminate design of the wing root joint is used for the entire wing structural box (upper and lower skin panels, ribs, and front and rear spars). The structure is purposely over designed to provide a starting point for subsequent structural optimization. The wing model will be made freely available to researchers.

Experimental and Computational Studies of Twin-Vertical-Tail Buffet

Abstract: Characteristics and mechanism of twin-vertical-tail buffet response on airplane configuration with wing root leading edge extension (LEX) were studied by both experiment and computation. Low-speed wind tunnel experiments were carried out to measure the root bending moment and tip acceleration of vertical tail. Vortical flow patterns were visualized via laser light sheet technique. Three-dimensional computation was performed to solve the unsteady Euler equations on rigid model. The results indicate that (1) bursting of vortices emanating from LEX is the main source of twin-vertical-tail buffet; (2) the Euler equations is able to predict the general characteristics of vertical-tail buffet response reasonably.